Rotor assembly disk spacer for a gas turbine engine

ABSTRACT

A disk spacer for a gas turbine engine includes a rim, a bore, a web and a load path spacer. The web extends between the rim and the bore. The load path spacer is positioned between the rim and the bore.

BACKGROUND OF THE INVENTION

This application relates generally to a gas turbine engine, and moreparticularly to a disk spacer for a rotor assembly of a gas turbineengine that provides anti-vortexing features.

A gas turbine engine channels airflow through its core along a desiredflow path. Many sections of the gas turbine engine must be cooled toensure reliable performance and efficiency. For example, the turbinesection of the gas turbine engine may require cooling airflow from acompressor section of the gas turbine engine. Cooling airflow of thistype may be provided by extracting bleed airflow from the core flow pathof the compressor section.

Anti-vortex tubes are used to provide radial inflow of bleed airflowextracted from the core flow path of a compressor section. Theanti-vortex tubes reduce vortices within the bleed airflow that reducethe radial inflow of the conditioned bleed airflow. The anti-vortextubes act as an impeller to raise the pressure of the bleed airflow andprepare the bleed airflow to cool portions of the downstream sections ofthe gas turbine engine.

The anti-vortex tubes are typically mounted to one or more rotor disksof a rotor assembly and therefore increase the weight carrying load ofsuch rotor disks. The increased weight of the rotor disks may reduce thefatigue life of such components.

SUMMARY

A disk spacer for a gas turbine engine includes a rim, a bore, a webthat extends between the rim and the bore, and a load path spacer. Theload path spacer is positioned between the rim and the bore.

In another exemplary embodiment, a gas turbine engine includes a rotorassembly having a first rotor disk, a second rotor disk, and a diskspacer. The disk spacer extends axially between the first rotor disk andthe second rotor disk. The disk spacer includes a rim, a bore, a web anda load path spacer. The web extends between the rim and the bore. Theload path spacer is positioned between the rim and the bore.

In another exemplary embodiment, a method of providing a rotor assemblyhaving a disk spacer for a gas turbine engine includes positioning thedisk spacer axially between a first rotor disk and a second rotor disk,communicating a bleed airflow from a core flow path of the gas turbineengine into a cavity of the gas turbine engine, increasing a pressure ofthe bleed airflow at a location between the first rotor disk and secondrotor disk within the cavity, and communicating the bleed airflow fromthe cavity to an axially downstream portion of the gas turbine engine.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a simplified cross-sectional view of a standard gasturbine engine;

FIG. 2 illustrates a cross-sectional view of a portion of the gasturbine engine; and

FIG. 3 illustrates an example rotor assembly that includes a diskspacer.

DETAILED DESCRIPTION

FIG. 1 shows a gas turbine engine 10, such as a turbo fan gas turbineengine, that is circumferentially disposed about an engine centerline(or axial centerline axis) 12. The gas turbine engine 10 includes a fansection 14, a compressor section 15 having a low pressure compressor 16and a high pressure compressor 18, a combustor 20 and a turbine section21 including a high pressure turbine 22 and a low pressure turbine 24.This application can also extend to engines without a fan, and with moreor fewer sections.

As is known, air is compressed in a low pressure compressor 16 and thehigh pressure compressor 18, is mixed with fuel and burned in thecombustor 20, and is expanded in the high pressure turbine 22 and a lowpressure turbine 24. Rotor assemblies 26 rotate in response to theexpansion, driving the low pressure and high pressure compressors 16, 18and the fan section 14. The low and high pressure compressors 16, 18include alternating rows of rotating compressor rotor airfoils or blades28 and static stator vanes 30. The high and low pressure turbines 22, 24include alternating rows of rotating turbine rotor airfoils or blades 32and static stator vanes 34.

It should be understood that this view is included simply to provide abasic understanding of the sections of a gas turbine engine 10 and notto limit the disclosure. This disclosure extends to all types of gasturbine engines 10 and for all types of applications.

FIG. 2 shows a portion of the compressor section 15 of a gas turbineengine 10. In this example, the portion depicted is the high pressurecompressor 18 of a gas turbine engine 10. However, this disclosure isnot limited to the high pressure compressor 18, and could extend toother sections of the gas turbine engine 10.

The illustrated compressor section 15 includes multiple stages ofalternating rows of rotor assemblies 26 and stator vanes 30. Althoughdepicted with a specific number of stages, the compressor section 15could include more or less stages. The stator vanes 30 extend betweeneach rotor assembly 26. Each rotor assembly 26 includes a rotor airfoil28 and a rotor disk 36. The rotor disks 36 include a rim 38, a bore 40,and a web 42 that extends between the rim 38 and the bore 40. A cavity44 extends between adjacent rotor disks 36.

A tie shaft 46 is connected to the rotor assemblies 26. The tie shaft 46can be preloaded to maintain tension on the plurality of rotorassemblies 26. The tie shaft 46 extends between a forward hub (notshown) and an aft hub 50. The tie shaft 46 may be threaded through theforward hub and snapped into the rotor disk 36 of the rotor assembly 26of the final stage of the compressor section 15. Once connected betweenthe forward hub and the aft hub 50, the preloaded tension on the tieshaft 46 is maintained with a nut 52.

At least one rotor assembly 26 includes a disk spacer 54 axiallypositioned between adjacent rotor disks 36. That is, the disk spacer 54is positioned within the cavity 44 that is defined between two adjacentrotor disks 36. The disk spacer 54 is axially offset from the rotorairfoils 28 of adjacent rotor disks 36. In one example, the disk spacer54 is positioned in a forward stage of the compressor section 15. Theactual location of the example disk spacer 54 will vary depending upon anumber of factors, including but not limited to, the desired systemlevel efficiency, the cooling flow requirements of the gas turbineengine components, bleed location requirements, and heat transferrequirements.

The example disk spacer includes a rim 56, a bore 58, a web 60 thatextends between the rim 56 and the bore 58, and a load path spacer 62that is positioned between the rim 56 and the bore 58. Among otherfunctions, the disk spacer 54 maintains a desired positioning andsupports a load of adjacent rotor disks 36, provides a flow path forreceiving a bleed airflow B that is extracted from a core flow path Ccommunicated through the compressor section 15, and providesanti-vortexing features. The bleed airflow B is communicated through thedisk spacer 54, into the cavity 44, under the bores 40 of the rotordisks 36, and aft to the downstream sections of the gas turbine engine10, such as the turbine section 21, for example.

FIG. 3 illustrates an example disk spacer 54 of a rotor assembly 26. Thedisk spacer 54 is radially trapped between a first rotor disk 36A and asecond rotor disk 36B. In this example, the load path spacer 62 of thedisk spacer 54 is radially trapped between the rim 38 of the rotor disk36A and the rim 38 of the rotor disk 36B. Friction forces between theload path spacer 62 and the rims 38 of the first rotor disk 36A and thesecond rotor disk 36B minimize any circumferential movement of the diskspacer 54 relative to the rotor disks 36. The load path spacer 62 caninclude a cylindrical shape, a conical shape, a catenary shape or anyother shape suitable for bridging the distance between adjacent rotordisks 36. The actual shape of the load path spacer 62 may vary dependingupon the size, shape and configuration of adjacent rotor disks 36.Because the disk spacer 54 is radially trapped between the first rotordisk 36A and the second rotor disk 36B, the disk spacer 54 will rotateas the rotor disks 36A, 36B rotate during operation. The disk spacer 54is positioned in a forward stage of the compressor section 15, in thisexample. However, other locations may be suitable for the disk spacer54.

The disk spacer 54 is generally cross-shaped. The rim 56 of the diskspacer 54 is positioned radially inwardly from a stator vane 30 of thecompressor section 15. In one example, the rim 56 is coaxial with aradial axis R of the stator vane 30. The rim 56 can also include a sealcoating 57, such as Zirconium Oxide, to seal the interface between thestator vane 30 and the rim 56 to reduce the potential for damage to thestator vane 30.

At least one gap 64 extends between an axially outermost tip 55 of therim 56 of the disk spacer 54 and the rim 38 of at least one of theadjacent rotor disks 36A, 36B. In this example, the gaps 64 extendbetween the axially outermost tips 55 of the rim 56 of the disk spacer54 and the rims 38 of each of the adjacent first rotor disk 36A andsecond rotor disk 36B. The gaps 64 provide a flow path for extractingbleed airflow B from the core flow path C of the compressor section 15.That is, the bleed airflow B is permitted to escape through the gaps 64toward the load path spacer 62 during operation.

The disk spacer 54 is defined along a longitudinal axis Al. The web 60of the disk spacer 54 extends substantially along the longitudinal axisAl. In this example, the load path spacer 62 extends along an axis A2that is generally transverse to the longitudinal axis Al of the diskspacer 54. The load path spacer 62 may extend generally parallel to therim 56 of the disk spacer 54, and in another example, may extendslightly transverse relative to the rim 56 of the disk spacer 54. Aportion 66 of the web 60 extends between the rim 56 and the load pathspacer 62 of the disk spacer 54. In this example, the portion 66 of theweb 60 is a smaller than the portion of the web 60 that extends betweenthe load path spacer 62 and the bore 58.

The first rotor disk 36A and the second rotor disk 36B are defined alonglongitudinal axes A3 and A4, respectively. In this example, thelongitudinal axis Al of the disk spacer 54 is substantially parallel tothe longitudinal axes A3, A4 of the first and second rotor disks 36A,36B.

The load path spacer 62 includes a plurality of openings 68 thatcommunicate the bleed airflow B through the disk spacer 54 and into thecavity 44 defined between adjacent rotor disks 36A and 36B. The openingsextend along axes A5 that are generally parallel to the longitudinalaxis A1 of the disk spacer 54. In this example, an axially upstream tipof the load path spacer 62 extends axially upstream relative to theaxially outermost tip 55 of the rim 56 and an axially downstream tip ofthe load path spacer 54 extends axially downstream relative to theaxially outermost tip 55 of the rim 56.

The web 60 of the disk spacer 54 includes fins 70. The fins 70 functionas an impeller on the bleed airflow B. For example, during operation,the disk spacer 54 rotates in unison with the rotor disks 36A, 36B.During rotation, the rotation of the fins 70 conditions the bleedairflow B and communicates the bleed airflow B radially inward(direction RI) toward the centerline axis 12 of the gas turbine engine10. That is, the fins 70 provide a radial inflow of the bleed airflow Bwithin the cavity 44. The conditioned bleed airflow B then travels belowthe bores 40 before exiting axially rearwardly for heat exchange withthe downstream components of the gas turbine engine 10.

The fins 70 of the web 60 further provide anti-vortexing features, suchas increasing the pressure of the bleed airflow B to prepare suchairflow to cool the downstream sections of the gas turbine engine 10.For example, once the pressure of the bleed airflow B is increased bythe fins 70 of the disk spacer 54, the bleed airflow B is communicatedaxially downstream to the high pressure turbine section 22 for coolingthereof.

In this example, the fins 70 extend along the web 60 of the disk spacer54 between the load path spacer 62 and the bore 58. In one example, thefins 70 are integrally formed with the web 60. In another example, thefins 70 are mechanically attached to the web 60.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claims should bestudied to determine the true scope and content of this disclosure.

1. A disk spacer for a gas turbine engine, comprising: a rim; a bore; aweb that extends between said rim and said bore, and a load path spacerpositioned between said rim and said bore.
 2. The disk spacer as recitedin claim 1, wherein said load path spacer includes openings operable tocommunicate a bleed airflow through said disk spacer.
 3. The disk spaceras recited in claim 1, wherein said web includes at least one fin. 4.The disk spacer as recited in claim 3, wherein said at least one finextends along said web between said load path spacer and said bore. 5.The disk spacer as recited in claim 1, wherein said web extends along alongitudinal axis, and said load path spacer extends along an axis thatis transverse to said longitudinal axis.
 6. The disk spacer as recitedin claim 1, wherein a portion of said web extends between said rim andsaid load path spacer.
 7. A rotor assembly for a gas turbine engine,comprising: a first rotor disk; a second rotor disk; and a disk spacerthat extends axially between said first rotor disk and said second rotordisk, wherein said disk spacer includes a rim, a bore, a web thatextends between said rim and said bore, and a load path spacerpositioned between said rim and said bore.
 8. The assembly as recited 7,wherein the rotor assembly is positioned within a compressor section ofthe gas turbine engine.
 9. The assembly as recited in claim 7, whereinsaid load path spacer is radially trapped between said first rotor diskand said second rotor disk.
 10. The assembly as recited 7, wherein saidload path spacer includes openings that communicate bleed airflow into acavity that extends between said first rotor disk and said second rotordisk.
 11. The assembly as recited in claim 7, wherein said web includesat least one fin.
 12. The assembly as recited in claim 11, wherein saidat least one fin extends between said load path spacer and said bore.13. The assembly as recited in claim 7, wherein said rim is positionedradially inwardly from a static stator vane of the gas turbine engine.14. The assembly as recited in claim 7, wherein a gap extends between anaxially outermost tip of said rim and each of said first rotor disk andsaid second rotor disk, and said gap defines a flow path for receiving ableed airflow from a core flow path of the gas turbine engine.
 15. Theassembly as recited in claim 7, wherein said first rotor disk extendsalong a first longitudinal axis, said second rotor disk extends along asecond longitudinal axis, and said disk spacer extends along a thirdlongitudinal axis, and said third longitudinal axis is parallel witheach of said first longitudinal axis and said second longitudinal axis.16. A method of providing a rotor assembly having a disk spacer for agas turbine engine, comprising the steps of: positioning the disk spaceraxially between a first rotor disk and a second rotor disk of the rotorassembly; communicating a bleed airflow from a core flow path of the gasturbine engine into a cavity of the gas turbine engine; increasing apressure of the bleed airflow at a location between the first rotor diskand the second rotor disk within the cavity; and communicating the bleedairflow from the cavity to an axially downstream portion of the gasturbine engine.
 17. The method as recited in claim 16, wherein the stepof positioning the disk spacer includes: radially trapping a load pathspacer of the disk spacer between the first rotor disk and the secondrotor disk.
 18. The method as recited in claim 16, wherein the step ofcommunicating the bleed airflow includes: providing openings through aportion of the disk spacer; and communicating the bleed airflow throughthe openings and into the cavity.
 19. The method as recited in claim 16,wherein the step of increasing the pressure of the bleed airflowincludes: increasing a radial inflow of the bleed airflow with at leastone fin.
 20. The method as recited in claim 16, wherein the step ofcommunicating the bleed airflow includes: communicating the bleedairflow from the cavity of a compressor section axially downstream to aturbine section of the gas turbine engine.